Gas turbine engine with third stream

ABSTRACT

A gas turbine is provided, the gas turbine engine including a turbomachine having an inlet splitter defining in part an inlet to a working gas flowpath and a fan duct splitter defining in part an inlet to a fan duct flowpath. The gas turbine engine also includes a primary fan driven by the turbomachine defining a primary fan tip radius R1, a primary fan hub radius R2, and a primary fan specific thrust rating TP; and a secondary fan downstream of the primary fan and driven by the turbomachine, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating TS; wherein the gas turbine engine defines an Effective Bypass Area, and wherein a ratio of R1 to R3 equals 
     
       
         
           
             
               
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FIELD

The present disclosure relates to a gas turbine engine with a thirdstream.

BACKGROUND

A gas turbine engine typically includes a fan and a turbomachine. Theturbomachine generally includes an inlet, one or more compressors, acombustor, and at least one turbine. The compressors compress air whichis channeled to the combustor where it is mixed with fuel. The mixtureis then ignited for generating hot combustion gases. The combustiongases are channeled to the turbine(s) which extracts energy from thecombustion gases for powering the compressor(s), as well as forproducing useful work to propel an aircraft in flight. The turbomachineis mechanically coupled to the fan for driving the fan during operation.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present disclosure, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of a three-stream engine inaccordance with an exemplary embodiment of the present disclosure.

FIG. 2 is a close-up, schematic view of the exemplary three-streamengine of FIG. 1 .

FIG. 3 is a close-up view of an area surrounding a leading edge of acore cowl of the exemplary three-stream engine of FIG. 2 .

FIG. 4 is a graph depicting exemplary primary fan specific thrust ratingT_(P) values and exemplary secondary fan specific thrust rating T_(S)values in accordance with exemplary aspects of the present disclosure.

FIGS. 5A through 5F is a table of example embodiments of the presentdisclosure.

FIGS. 6A through 6C are graphs depicting a range of radius ratios toEffective Bypass Areas in accordance with various example embodiments ofthe present disclosure.

FIG. 7 is a schematic view of a turboprop engine in accordance with anexemplary aspect of the present disclosure.

FIG. 8 is a schematic view of a direct drive, ducted, turbofan engine inaccordance with an exemplary aspect of the present disclosure.

FIG. 9 is a schematic view of a geared, ducted, turbofan engine inaccordance with an exemplary aspect of the present disclosure.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of thedisclosure, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the disclosure.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations. Additionally, unlessspecifically identified otherwise, all embodiments described hereinshould be considered exemplary.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

A “third stream” as used herein means a non-primary air stream capableof increasing fluid energy to produce a minority of total propulsionsystem thrust. A pressure ratio of the third stream may be higher thanthat of the primary propulsion stream (e.g., a bypass or propellerdriven propulsion stream). The thrust may be produced through adedicated nozzle or through mixing of an airflow through the thirdstream with a primary propulsion stream or a core air stream, e.g., intoa common nozzle.

In certain exemplary embodiments an operating temperature of the airflowthrough the third stream may be less than a maximum compressor dischargetemperature for the engine, and more specifically may be less than 350degrees Fahrenheit (such as less than 300 degrees Fahrenheit, such asless than 250 degrees Fahrenheit, such as less than 200 degreesFahrenheit, and at least as great as an ambient temperature). In certainexemplary embodiments these operating temperatures may facilitate heattransfer to or from the airflow through the third stream and a separatefluid stream. Further, in certain exemplary embodiments, the airflowthrough the third stream may contribute less than 50% of the totalengine thrust (and at least, e.g., 2% of the total engine thrust) at atakeoff condition, or more particularly while operating at a ratedtakeoff power at sea level, static flight speed, 86 degrees Fahrenheitambient temperature operating conditions.

Furthermore in certain exemplary embodiments, aspects of the airflowthrough the third stream (e.g., airstream, mixing, or exhaustproperties), and thereby the aforementioned exemplary percentcontribution to total thrust, may passively adjust during engineoperation or be modified purposefully through use of engine controlfeatures (such as fuel flow, electric machine power, variable stators,variable inlet guide vanes, valves, variable exhaust geometry, orfluidic features) to adjust or optimize overall system performanceacross a broad range of potential operating conditions.

The term “disk loading” refers to an average pressure change across aplurality of rotor blades of a rotor assembly, such as the averagepressure change across a plurality of fan blades of a fan.

The term “rated speed” refers to an operating condition of an enginewhereby the engine is operating in the maximum, full load operatingcondition that is rated by the manufacturer.

The term “standard day operating condition” refers to ambient conditionsof sea level altitude, 59 degrees Fahrenheit, and 60 percent relativehumidity.

The term “propulsive efficiency” refers to an efficiency with which theenergy contained in an engine's fuel is converted into kinetic energyfor the vehicle incorporating the engine, to accelerate it, or toreplace losses due to aerodynamic drag or gravity.

The term “bypass ratio” refers to a ratio in an engine of an amount ofairflow that is bypassed around the engine's ducted inlet to the amountthat passes through the engine's ducted inlet. For example, in theembodiment of FIG. 1 , discussed below, the bypass ratio refers to anamount of airflow from the fan 152 that flows over the fan cowl 170 toan amount of airflow from the fan 152 that flows through the engineinlet 182.

The term “corrected tip speed,” with respect to a fan having fan blades,refers to a speed of the fan blades at an outer tip of the fan bladesalong a radial direction, corrected to correspond to standard dayconditions (i.e., the speed the fan blades at their outer tips wouldrotate at if the upstream temperature corresponded to standard dayconditions). A corrected tip speed of a rotor (such as the fan) may becalculated by dividing a physical speed by the square root of an averagerotor inlet temperature (in Rankine) over a reference temperature of518.67 Rankine.

Generally, a turbofan engine includes a relatively large fan to providea desired amount of thrust without overloading the fan blades (i.e.,without increasing a disk loading of the fan blades of the fan beyond acertain threshold), and therefore to maintain a desired overallpropulsive efficiency for the turbofan engine. Conventional turbofanengine design practice has been to provide a large fan, or rather a highdiameter fan, on the engine to provide as much of a total thrust for theturbofan engine as reasonably possible. The objective, when designingthe conventional turbofan engine was to maximize a propulsive efficiencyof the turbofan engine. A turbofan engine including such a large fan,however, may result in, e.g., problems packaging the turbofan engine onan aircraft, a relatively heavy turbofan engine (particularly for ductedturbofan engines), etc. Further, as the need for turbofan engines toprovide more thrust continues, the thermal demands on the turbofanengines correspondingly increases.

The inventors of the present disclosure, however, found that for a threestream turbofan engine having a primary fan and a secondary fan, withthe secondary fan being a ducted fan providing an airflow to a thirdstream of the engine, the amount of thrust generation required from theprimary fan may be reduced, with the secondary fan providing thedifference through the third stream. Such a configuration may maintain adesired overall propulsive efficiently for the turbofan engine, orunexpectedly may in fact increase the over propulsive efficiency of theturbofan engine.

The inventors proceeded in the manner of designing an engine with givenprimary fan characteristics, secondary fan characteristics, andturbomachine characteristics; checking the propulsive efficiency of thedesigned turbofan engine; redesigning the turbofan engine with varyingprimary fan, secondary fan, and turbomachine characteristics; recheckingthe propulsive efficiency of the redesigned turbofan engine; etc. duringthe design of several different types of turbofan engines, including thegas turbine engine described below with reference to FIGS. 1, 2, 7, 8and 9 . During the course of this practice of studying/evaluatingvarious primary fan characteristics, secondary fan characteristics, andturbomachine characteristics considered feasible for best satisfyingmission requirements, it was discovered that a certain relationshipexists between a percentage of a total turbofan engine thrust providedby a third stream (as defined herein) and the relative sizes of aturbofan's primary to secondary fan, or more particularly a radius ratioof the primary fan to secondary fan.

Moreover, it was discovered that in lieu of calculating the percentageof total turbofan engine thrust provided by the third stream for eachdesign, which may be difficult to accurately calculate across variousoperating conditions, ambient conditions, engine designs, etc., anarea-weighted specific thrust rating for the primary fan and for thesecondary fan (area weighted based on an inlet area to the third streamand an inlet area to a bypass passage) may effectively and accuratelyrepresent the percentage of total turbofan engine thrust provided by thethird stream.

The resulting radius ratio to third-stream thrust relationship, asherein referred to, can be thought of as an indicator of the ability ofa turbofan engine to maintain or even improve upon a desired propulsiveefficiency via the third stream and, additionally, indicating animprovement in the turbofan engine's packaging concerns and weightconcerns, and thermal management capabilities.

Referring now to FIG. 1 , a schematic cross-sectional view of a gasturbine engine 100 is provided according to an example embodiment of thepresent disclosure. Particularly, FIG. 1 provides a turbofan enginehaving a rotor assembly with a single stage of unducted rotor blades. Insuch a manner, the rotor assembly may be referred to herein as an“unducted fan,” or the entire engine 100 may be referred to as an“unducted turbofan engine.” In addition, the engine 100 of FIG. 1includes a third stream extending from the compressor section to a rotorassembly flowpath over the turbomachine, as will be explained in moredetail below.

For reference, the engine 100 defines an axial direction A, a radialdirection R, and a circumferential direction C. Moreover, the engine 100defines an axial centerline or longitudinal axis 112 that extends alongthe axial direction A. In general, the axial direction A extendsparallel to the longitudinal axis 112, the radial direction R extendsoutward from and inward to the longitudinal axis 112 in a directionorthogonal to the axial direction A, and the circumferential directionextends three hundred sixty degrees (360°) around the longitudinal axis112. The engine 100 extends between a forward end 114 and an aft end116, e.g., along the axial direction A.

The engine 100 includes a turbomachine 120 and a rotor assembly, alsoreferred to a fan section 150, positioned upstream thereof. Generally,the turbomachine 120 includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. Particularly, as shown in FIG. 1 , the turbomachine 120includes a core cowl 122 that defines an annular core inlet 124. Thecore cowl 122 further encloses at least in part a low pressure systemand a high pressure system. For example, the core cowl 122 depictedencloses and supports at least in part a booster or low pressure (“LP”)compressor 126 for pressurizing the air that enters the turbomachine 120through core inlet 124. A high pressure (“HP”), multi-stage, axial-flowcompressor 128 receives pressurized air from the LP compressor 126 andfurther increases the pressure of the air. The pressurized air streamflows downstream to a combustor 130 of the combustion section where fuelis injected into the pressurized air stream and ignited to raise thetemperature and energy level of the pressurized air.

It will be appreciated that as used herein, the terms “high/low speed”and “high/low pressure” are used with respect to the high pressure/highspeed system and low pressure/low speed system interchangeably. Further,it will be appreciated that the terms “high” and “low” are used in thissame context to distinguish the two systems, and are not meant to implyany absolute speed and/or pressure values.

The high energy combustion products flow from the combustor 130downstream to a high pressure turbine 132. The high pressure turbine 128drives the high pressure compressor 128 through a high pressure shaft136. In this regard, the high pressure turbine 128 is drivingly coupledwith the high pressure compressor 128. The high energy combustionproducts then flow to a low pressure turbine 134. The low pressureturbine 134 drives the low pressure compressor 126 and components of thefan section 150 through a low pressure shaft 138. In this regard, thelow pressure turbine 134 is drivingly coupled with the low pressurecompressor 126 and components of the fan section 150. The LP shaft 138is coaxial with the HP shaft 136 in this example embodiment. Afterdriving each of the turbines 132, 134, the combustion products exit theturbomachine 120 through a turbomachine exhaust nozzle 140.

Accordingly, the turbomachine 120 defines a working gas flowpath or coreduct 142 that extends between the core inlet 124 and the turbomachineexhaust nozzle 140. The core duct 142 is an annular duct positionedgenerally inward of the core cowl 122 along the radial direction R. Thecore duct 142 (e.g., the working gas flowpath through the turbomachine120) may be referred to as a second stream.

The fan section 150 includes a fan 152, which is the primary fan in thisexample embodiment. For the depicted embodiment of FIG. 1 , the fan 152is an open rotor or unducted fan 152. In such a manner, the engine 100may be referred to as an open rotor engine.

As depicted, the fan 152 includes an array of fan blades 154 (only oneshown in FIG. 1 ). The fan blades 154 are rotatable, e.g., about thelongitudinal axis 112. As noted above, the fan 152 is drivingly coupledwith the low pressure turbine 134 via the LP shaft 138. For theembodiments shown in FIG. 1 , the fan 152 is coupled with the LP shaft138 via a speed reduction gearbox 155, e.g., in an indirect-drive orgeared-drive configuration.

Moreover, the array of fan blades 154 can be arranged in equal spacingaround the longitudinal axis 112. Each fan blade 154 has a root and atip and a span defined therebetween. Further, each fan blade 154 definesa fan blade tip radius R₁ along the radial direction R from thelongitudinal axis 112 to the tip, and a hub radius (or inner radius) R₂along the radial direction R from the longitudinal axis 112 to the base.Further, the fan 152, or rather each fan blade 154 of the fan 152,defines a fan radius ratio, RqR, equal to R₁ divided by R₂. As the fan152 is the primary fan of the engine 100, the fan radius ratio, RqR, ofthe fan 152 may be referred to as the primary fan radius ratio,RqR_(Prim.-Fan).

Moreover, each fan blade 154 defines a central blade axis 156. For thisembodiment, each fan blade 154 of the fan 152 is rotatable about theirrespective central blade axis 156, e.g., in unison with one another. Oneor more actuators 158 are provided to facilitate such rotation andtherefore may be used to change a pitch of the fan blades 154 abouttheir respective central blades' axes 156.

The fan section 150 further includes a fan guide vane array 160 thatincludes fan guide vanes 162 (only one shown in FIG. 1 ) disposed aroundthe longitudinal axis 112. For this embodiment, the fan guide vanes 162are not rotatable about the longitudinal axis 112. Each fan guide vane162 has a root and a tip and a span defined therebetween. The fan guidevanes 162 may be unshrouded as shown in FIG. 1 or, alternatively, may beshrouded, e.g., by an annular shroud spaced outward from the tips of thefan guide vanes 162 along the radial direction R or attached to the fanguide vanes 162.

Each fan guide vane 162 defines a central blade axis 164. For thisembodiment, each fan guide vane 162 of the fan guide vane array 160 isrotatable about its respective central blade axis 164, e.g., in unisonwith one another. One or more actuators 166 are provided to facilitatesuch rotation and therefore may be used to change a pitch of the fanguide vane 162 about its respective central blade axis 164. However, inother embodiments, each fan guide vane 162 may be fixed or unable to bepitched about its central blade axis 164. The fan guide vanes 162 aremounted to the fan cowl 170.

As shown in FIG. 1 , in addition to the fan 152, which is unducted, aducted fan 184 is included aft of the fan 152, such that the engine 100includes both a ducted and an unducted fan which both serve to generatethrust through the movement of air without passage through at least aportion of the turbomachine 120 (e.g., without passage through the HPcompressor 128 and combustion section for the embodiment depicted). Theducted fan 184 is rotatable about the same axis (e.g., the longitudinalaxis 112) as the fan blade 154. The ducted fan 184 is, for theembodiment depicted, driven by the low pressure turbine 134 (e.g.coupled to the LP shaft 138). In the embodiment depicted, as notedabove, the fan 152 may be referred to as the primary fan, and the ductedfan 184 may be referred to as a secondary fan. It will be appreciatedthat these terms “primary” and “secondary” are terms of convenience, anddo not imply any particular importance, power, or the like.

The ducted fan 184 includes a plurality of fan blades (not separatelylabeled in FIG. 1 ; see fan blades 185 labeled in FIG. 2 ) arranged in asingle stage, such that the ducted fan 184 may be referred to as asingle stage fan. The fan blades of the ducted fan 184 can be arrangedin equal spacing around the longitudinal axis 112. Each blade of theducted fan 184 has a root and a tip and a span defined therebetween.Further, each fan blade of the ducted fan 184 defines a fan blade tipradius R₃ along the radial direction R from the longitudinal axis 112 tothe tip, and a hub radius (or inner radius) R₄ along the radialdirection R from the longitudinal axis 112 to the base. Further, theducted fan 184, or rather each fan blade of the ducted fan 184, definesa fan radius ratio, RqR, equal to R₃ divided by R₄. As the ducted fan184 is the secondary fan of the engine 100, the fan radius ratio, RqR,of the ducted fan 184 may be referred to as the secondary fan radiusratio, RqR_(Sec.-Fan).

The fan cowl 170 annularly encases at least a portion of the core cowl122 and is generally positioned outward of at least a portion of thecore cowl 122 along the radial direction R. Particularly, a downstreamsection of the fan cowl 170 extends over a forward portion of the corecowl 122 to define a fan duct flowpath, or simply a fan duct 172.According to this embodiment, the fan flowpath or fan duct 172 may beunderstood as forming at least a portion of the third stream of theengine 100.

Incoming air may enter through the fan duct 172 through a fan duct inlet176 and may exit through a fan exhaust nozzle 178 to produce propulsivethrust. The fan duct 172 is an annular duct positioned generally outwardof the core duct 142 along the radial direction R. The fan cowl 170 andthe core cowl 122 are connected together and supported by a plurality ofsubstantially radially-extending, circumferentially-spaced stationarystruts 174 (only one shown in FIG. 1 ). The stationary struts 174 mayeach be aerodynamically contoured to direct air flowing thereby. Otherstruts in addition to the stationary struts 174 may be used to connectand support the fan cowl 170 and/or core cowl 122. In many embodiments,the fan duct 172 and the core duct 142 may at least partially co-extend(generally axially) on opposite sides (e.g., opposite radial sides) ofthe core cowl 122. For example, the fan duct 172 and the core duct 142may each extend directly from a leading edge 144 of the core cowl 122and may partially co-extend generally axially on opposite radial sidesof the core cowl.

The engine 100 also defines or includes an inlet duct 180. The inletduct 180 extends between the engine inlet 182 and the core inlet 124/fanduct inlet 176. The engine inlet 182 is defined generally at the forwardend of the fan cowl 170 and is positioned between the fan 152 and thefan guide vane array 160 along the axial direction A. The inlet duct 180is an annular duct that is positioned inward of the fan cowl 170 alongthe radial direction R. Air flowing downstream along the inlet duct 180is split, not necessarily evenly, into the core duct 142 and the fanduct 172 by a fan duct splitter or leading edge 144 of the core cowl122. The inlet duct 180 is wider than the core duct 142 along the radialdirection R. The inlet duct 180 is also wider than the fan duct 172along the radial direction R.

Notably, for the embodiment depicted, the engine 100 includes one ormore features to increase an efficiency of a third stream thrust,Fn_(3S) (e.g., a thrust generated by an airflow through the fan duct 172exiting through the fan exhaust nozzle 178, generated at least in partby the ducted fan 184). In particular, the engine 100 further includesan array of inlet guide vanes 186 positioned in the inlet duct 180upstream of the ducted fan 184 and downstream of the engine inlet 182.The array of inlet guide vanes 186 are arranged around the longitudinalaxis 112. For this embodiment, the inlet guide vanes 186 are notrotatable about the longitudinal axis 112. Each inlet guide vanes 186defines a central blade axis (not labeled for clarity), and is rotatableabout its respective central blade axis, e.g., in unison with oneanother. In such a manner, the inlet guide vanes 186 may be considered avariable geometry component. One or more actuators 188 are provided tofacilitate such rotation and therefore may be used to change a pitch ofthe inlet guide vanes 186 about their respective central blade axes.However, in other embodiments, each inlet guide vanes 186 may be fixedor unable to be pitched about its central blade axis.

Further, located downstream of the ducted fan 184 and upstream of thefan duct inlet 176, the engine 100 includes an array of outlet guidevanes 190. As with the array of inlet guide vanes 186, the array ofoutlet guide vanes 190 are not rotatable about the longitudinal axis112. However, for the embodiment depicted, unlike the array of inletguide vanes 186, the array of outlet guide vanes 190 are configured asfixed-pitch outlet guide vanes.

Further, it will be appreciated that for the embodiment depicted, thefan exhaust nozzle 178 of the fan duct 172 is further configured as avariable geometry exhaust nozzle. In such a manner, the engine 100includes one or more actuators 192 for modulating the variable geometryexhaust nozzle. For example, the variable geometry exhaust nozzle may beconfigured to vary a total cross-sectional area (e.g., an area of thenozzle in a plane perpendicular to the longitudinal axis 112) tomodulate an amount of thrust generated based on one or more engineoperating conditions (e.g., temperature, pressure, mass flowrate, etc.of an airflow through the fan duct 172). A fixed geometry exhaust nozzlemay also be adopted.

The combination of the array of inlet guide vanes 186 located upstreamof the ducted fan 184, the array of outlet guide vanes 190 locateddownstream of the ducted fan 184, and the fan exhaust nozzle 178 mayresult in a more efficient generation of third stream thrust, Fn_(3S),during one or more engine operating conditions. Further, by introducinga variability in the geometry of the inlet guide vanes 186 and the fanexhaust nozzle 178, the engine 100 may be capable of generating moreefficient third stream thrust, Fn_(3S), across a relatively wide arrayof engine operating conditions, including takeoff and climb (where amaximum total engine thrust Fn_(Total), is generally needed) as well ascruise (where a lesser amount of total engine thrust, Fn_(Total), isgenerally needed).

Moreover, referring still to FIG. 1 , in exemplary embodiments, airpassing through the fan duct 172 may be relatively cooler (e.g., lowertemperature) than one or more fluids utilized in the turbomachine 120.In this way, one or more heat exchangers 200 may be positioned inthermal communication with the fan duct 172. For example, one or moreheat exchangers 200 may be disposed within the fan duct 172 and utilizedto cool one or more fluids from the core engine with the air passingthrough the fan duct 172, as a resource for removing heat from a fluid,e.g., compressor bleed air, oil or fuel.

Although not depicted, the heat exchanger 200 may be an annular heatexchanger extending substantially 360 degrees in the fan duct 172 (e.g.,at least 300 degrees, such as at least 330 degrees). In such a manner,the heat exchanger 200 may effectively utilize the air passing throughthe fan duct 172 to cool one or more systems of the engine 100 (e.g.,lubrication oil systems, compressor bleed air, electrical components,etc.). The heat exchanger 200 uses the air passing through duct 172 as aheat sink and correspondingly increases the temperature of the airdownstream of the heat exchanger 200 and exiting the fan exhaust nozzle178.

During operation of the engine 100 at an operating condition, the engine100 generates a total thrust, Fn_(Total). The operating condition may beoperation of the engine 100 at a rated speed during standard dayoperating condition. The total thrust is a sum of a first stream thrust,Fn_(1S) (e.g., a primary fan thrust generated by an airflow over the fancowl 170 and core cowl 122, generated by the fan 152, through a bypasspassage 194), the third stream thrust, Fn_(3S), and a second streamthrust, Fn_(2S) (e.g., a thrust generated by an airflow through the coreduct 142 exiting through the turbomachine exhaust nozzle 140).

It will be appreciated that, as noted briefly above, it may be difficultto accurately calculate a percentage of the total thrust, Fn_(Total),provided by the third stream, or rather the airflow through the fan duct172 exiting through the fan exhaust nozzle 178, across various operatingconditions, ambient conditions, engine designs, etc. The inventors havefound, however, that area-weighting a specific thrust rating for theprimary fan (a primary fan specific thrust rating T_(P)) and a specificthrust rating for the secondary fan (a secondary fan specific thrustrating T_(S)) provides an accurate representation of an expectedpercentage of the total thrust, Fn_(Total), provided by the thirdstream.

More specifically, referring to FIG. 2 , a close-up, simplified,schematic view of the turbofan engine 100 of FIG. 1 is provided. Theturbofan engine 100, as noted above includes a primary fan, or ratherfan 152 having fan blades 154, and a secondary fan, or rather ducted fan184 having fan blades 185. Airflow from the fan 152 is split between thebypass passage 194 and the inlet duct 180 by an inlet splitter 196.Airflow from the ducted fan 184 is split between the fan duct 172 andthe core duct 142 by the leading edge 144 (sometimes also referred to asthe fan duct splitter).

The exemplary turbofan engine 100 depicted in FIG. 2 further defines aprimary fan outer fan area, A_(P_Out), a primary fan inner fan area,A_(P_In), a secondary fan outer fan area, A_(S_Out), and a secondary faninner fan area, A_(S_In).

The primary fan outer fan area, A_(P_Out), refers to an area defined byan annulus representing a portion of the fan 152 located outward of theinlet splitter 196 of the fan cowl 170. In particular, the turbofanengine 100 further defines a fan cowl splitter radius, R₅. The fan cowlsplitter radius, R₅, is defined along the radial direction R from thelongitudinal axis 112 to the inlet splitter 196. The primary fan outerfan area refers to an area defined by the formula: πR₁ ²−πR₅ ².

The primary fan inner fan area, A_(P_In), refers to an area defined byan annulus representing a portion of the fan 152 located inward of theinlet splitter 196 of the fan cowl 170. In particular, the turbofanengine 100 further defines an engine inlet inner radius, R₆. The engineinlet inner radius, R₆, is defined along the radial direction R from thelongitudinal axis 112 to an inner casing defining the engine inlet 182directly inward along the radial direction R from the inlet splitter196. The primary fan inner fan area refers to an area defined by theformula: πR₅ ²−πR₆ ².

The secondary fan outer fan area, A_(S_Out), refers to an arearepresenting a portion an airflow from the ducted fan 184 that isprovided to the fan duct 172. In particular, the leading edge 144defines a leading edge radius, R₇, and the turbofan engine 100 definesan effective fan duct inlet outer radius, R₈ (see FIG. 3 ). The leadingedge radius, R₇, is defined along the radial direction R from thelongitudinal axis 112 to the leading edge 144.

Referring briefly to FIG. 3 , providing a close-up view of an areasurrounding the leading edge 144, the fan duct 172 defines a cross-wiseheight 198 measured from the leading edge 144 to the fan cowl 170 in adirection perpendicular to a mean flow direction 204 of an airflowthrough a forward 10% of the fan duct 172. Notably, an angle 206 isdefined by the mean flow direction 204 relative to a reference line 208extending parallel to the longitudinal axis 112. The angle 206 isreferred to as θ. In certain embodiments, the angle 206 may be between 5degrees and 80 degrees, such as between 10 degrees and 60 degrees. Theeffective fan duct inlet outer radius, R₈, is defined along the radialdirection R from the longitudinal axis 112 to where the cross-wiseheight 198 meets the fan cowl 170. The secondary fan outer fan area,A_(S_Out), refers to an area defined by the formula:

$\frac{\pi\left( {R_{8}^{2} - R_{7}^{2}} \right)}{\cos(\theta)}.$

Referring back to FIG. 2 , the secondary fan inner fan area, A_(S_In),refers to an area defined by an annulus representing a portion of theducted fan 184 located inward of the leading edge 144 of the core cowl122. In particular, the turbofan engine 100 further defines a core inletinner radius, R₉. The core inlet inner radius, R₉, is defined along theradial direction R from the longitudinal axis 112 to an inner casingdefining the core inlet 124 directly inward along the radial direction Rfrom the leading edge 144. The primary fan inner fan area refers to anarea defined by the formula: πR₇ ²−πR₉ ².

These areas, and in particular the primary fan outer fan area,A_(P_Out), and secondary fan outer fan area, A_(S_Out), may be used toweight the primary fan specific thrust rating T_(P) and secondary fanspecific thrust rating T_(S) to accurately represent the expectedpercentage of the total thrust, Fn_(Total), provided by the thirdstream, as will be explained in more detail below. Notably, as usedherein, a ratio of the primary fan outer fan area, A_(P_Out), to thesecondary fan outer fan area, A_(S_Out), is referred to herein as anEffective Bypass Area, or EBA.

As alluded earlier, the inventors discovered, unexpectedly during thecourse of gas turbine engine design—i.e., designing gas turbine engines(e.g., both ducted and unducted turbofan engines and turboprop engines)having a variety of different primary fan and secondary fancharacteristics, both physical and operational—and evaluating an overallpropulsive efficiency, a significant relationship exists between apercentage of a total gas turbine engine thrust provided by a thirdstream (as defined herein) and the relative sizes of a gas turbineengine's primary to secondary fan. The resulting radius ratio tothird-stream thrust relationship, as herein referred to, can be thoughtof as an indicator of the ability of a gas turbine engine to maintain oreven improve upon a desired propulsive efficiency via the third streamand, additionally, indicating an improvement in the gas turbine engine'spackaging concerns and weight concerns, and thermal managementcapabilities.

As will be appreciated, higher and lower third stream thrusts change thepackaging abilities of the gas turbine engine and the thermal sinkcapabilities of the gas turbine engine. For example, increased thrustfrom an airflow through the third stream generally means more airflow(on a mass flowrate basis) through the third stream, which in turn meanmore thermal capacity for such airflow. Further, the inventors foundthat if you provide too little thrust from the third stream, the gasturbine engine may be unnecessarily large (and thus more difficult topackage) and heavy, and further may not provide a desired amount ofthermal sink capabilities. If you provide too much thrust through thethird stream, the engine may not fully take advantage of relativelyefficient thrust that may be generated by the primary fan.

The above relationship may be a function of a bypass ratio of the gasturbine engine, which may generally be limited by reasonable enginetemperatures, including operating temperatures, such as exhaust gastemperatures (EGT). For example, as will be appreciated in view of theforegoing teaching, a radius of the primary fan relative to a radius ofthe secondary fan, along with a percentage of a total gas turbine enginethrust generated by an airflow through the third stream duringoperation, are each, in part, a function of the bypass ratio andtogether characterize the balancing in the relationship noted above.

Many aspects of an architecture dictate the bypass ratio of a gasturbine engine. For example, the bypass ratio is, in part, a function ofa corrected tip speed of the primary fan relative to a corrected tipspeed the secondary fan, as well as a specific thrust of the respectiveprimary and secondary fans. The specific thrusts of the primary andsecondary fans, in turn, are a function of a pressure ratio of theprimary and secondary fans, respectively, and a disk loading (alsoreferred to as a power loading) on the primary and secondary fans,respectively. These factors also affect the balancing in therelationship noted above, as will described in more detail below withreference to an effective fan parameter, EFP.

As noted above, the inventors of the present disclosure discovered arelationship between the percentage of engine thrust configured to beprovided by the airflow through the third stream and the radius ratio ofthe primary fan and secondary fan that can result in a gas turbineengine maintaining or even improving upon a desired propulsiveefficiency, while also taking into account the gas turbine engine'spackaging concerns and weight concerns, and also providing desiredthermal management capabilities. The relationship discovered, infra, canidentify an improved engine configuration suited for a particularmission requirement, one that takes into account installation, packagingand loading, thermal sink needs and other factors influencing theoptimal choice for an engine configuration.

In addition to yielding an improved gas turbine engine, as explained indetail above, utilizing this relationship, the inventors found that thenumber of suitable or feasible gas turbine engine designs incorporatinga primary fan and a secondary fan, and defining a third stream, capableof meeting both the propulsive efficiency requirements and packaging,weight, and thermal sink requirements could be greatly diminished,thereby facilitating a more rapid down selection of designs to consideras a gas turbine engine is being developed. Such a benefit provides moreinsight to the requirements for a given gas turbine engine well beforespecific technologies, integration and system requirements are developedfully. Such a benefit avoids late-stage redesign.

The desired relationship providing for the improved gas turbine engine,discovered by the inventors, is expressed as:

$\begin{matrix}{\frac{R_{1}}{R_{3}} = \sqrt{\left( {EFP} \right)\frac{\left( {1 - {RqR}_{{Sec}.{- {Fan}}}^{2}} \right)}{\left( {1 - {RqR}_{{Prim}.{- {Fan}}}^{2}} \right)}\left( \frac{T_{P}}{T_{S}} \right)\left( {EBA} \right)}} & (1)\end{matrix}$

where R₁ is a tip radius of the primary fan, R₂ is a hub radius of theprimary fan, R₃ is a tip radius of the secondary fan, R₄ is a hub radiusof the secondary fan, RqR_(Prim.-Fan) is the ratio of R₁ to R₂,RqR_(Sec.-Fan) is the ratio of R₃ to R₄, EFP is an effective fanparameter, T_(P) is a primary fan specific thrust rating, T_(S) is asecondary fan specific thrust rating, and EBA is the effective bypassarea.

EBA relates a primary fan outer fan area, A_(P_Out), to a secondary fanouter fan area, A_(S_Out), and is expressed as a percentage. Inparticular, EBA is represented by the following ratio:A_(P_Out)/A_(S_Out). A higher EBA corresponds to a larger primary fanbypass ratio and a low secondary fan bypass ratio. Conversely, a lowerEBA corresponds to a smaller primary fan bypass ratio and a highersecondary fan bypass ratio.

EFP is a function of a corrected tip speed of the primary fan, acorrected tip speed of the secondary fan, a disk loading of the primaryfan, and a disk loading of the secondary fan. EFP, by taking intoaccount the corrected tip speeds of the primary and secondary fans,accounts for such factors as the specific engine configuration (e.g.,geared, direct drive, etc.), which may have some influence on therelationship between tip radius ratio (R₁ to R₃) and the percent thrustthrough the third stream (% Fn_(3S)) for a turbofan engine having adesired propulsive efficiency. The relationship of these contributingfactors to EFP to the tip radius ratio (R₁ to R₃) and the percent thrustthrough the third stream (% Fn_(3S)) for a turbofan engine is describedin more detail above.

The primary fan specific thrust rating T_(P) and the secondary fanspecific thrust rating T_(S) reflect the corrected tip speeds for theprimary fan (fan 152 in the embodiment of FIGS. 1 and 2 ) and for thesecondary fan (ducted fan 184 in the embodiment of FIGS. 1 and 2 ) at arated speed during standard day operating conditions. Example values forthe primary fan specific thrust rating T_(P) and the secondary fanspecific thrust rating T_(S) as may be incorporated into Expression (1)are provided below and in FIG. 4 .

In particular, FIG. 4 provides a graph 300 showing exemplary primary fanspecific thrust ratings T_(P) along a Y-axis 302 and exemplary secondaryfan specific thrust ratings T_(S) along an X-axis 304. Notably, thegraph 300 has plotted example primary fan specific thrust ratings T_(P)and secondary fan specific thrust ratings T_(S) from at least certain ofthe example engines described below with reference to FIGS. 5A through5F. As will be appreciated from the examples plotted in FIG. 4 , thesecondary fan specific thrust rating T_(S) may be considered asgenerally higher than the primary fan specific thrust rating T_(P) for agiven engine, as the secondary fan is downstream of, and builds off of aflow from, the primary fan.

For the embodiments of FIG. 4 , the primary fan specific thrust ratingT_(P) ranges from 0.08 to 0.59, and in at least certain exemplaryaspects from 0.1 to 0.5. Similarly, the secondary fan specific thrustrating T_(S) ranges from 0.21 to 0.6, and in at least certain exemplaryaspects from 0.26 to 0.5. These broader ranges correspond to an area 306in FIG. 4 , and the narrower ranges correspond to an area 308 in FIG. 4

In particular, the graph 300 in FIG. 4 identifies four subregions of theprimary fan specific thrust rating T_(P) and the secondary fan specificthrust rating T_(S) values. In particular, the graph 300 of FIG. 4identifies a first subregion 310, a second subregion 312, a thirdsubregion 314, and a fourth subregion 316.

The first subregion 310 includes two example engines plotted—example 318and example 320. Example 318 corresponds to a turboprop engine (see,e.g., exemplary turboprop engine 526 described below with reference toFIG. 7 ) and example 320 corresponds to an open rotor engine, such asthe exemplary engine 100 of FIGS. 1 and 2 . Each of these engines is ageared engine (i.e., includes a gearbox between a driving shaft and afan shaft; see, e.g., gearbox 155 of FIG. 1 ) and further is an unductedengine (i.e., does not include an outer nacelle surrounding the primaryfan). The primary fan specific thrust rating T_(P) for the firstsubregion 310 ranges from 0.1 to 0.23 and the secondary fan specificthrust rating T_(S) for the first subregion 310 ranges from 0.35 to0.45.

It will be appreciated that the primary fan specific thrust ratingsT_(P) of the first subregion 310 are relatively low as compared to therest of the examples plotted in FIG. 4 for at least the reason that theengines are unducted engines, allowing for a larger fan diameter. With alarger diameter fan, a lower primary fan specific thrust ratings T_(P)is acceptable while still providing a desired amount of thrust for theengine (as the engine may rotate more slowly, with less thrust per unitarea). In such a manner, it will be appreciated that the primary fans ofthe examples 318, 320 may be relatively efficient as they will generallydefine a relatively small pressure ratio during operation.

Further, as noted, the engines plotted as examples 318, 320 are gearedengines and may define a relatively high gear ratio (e.g., as comparedto at least certain examples in the second subregion 312, discussedbelow). Nonetheless, the secondary fan specific thrust ratings T_(S) maystill be smaller (e.g., as compared to at least certain examples in thesecond subregion 312). In particular, since the secondary fan specificthrust rating T_(S) of the secondary fan builds off of the primary fanspecific thrust ratings T_(P) of the primary fan (and the primary fanspecific thrust ratings T_(P) are smaller as compared to the examples inthe second subregion 312), the secondary fan specific thrust ratingsT_(S) are also smaller.

The second subregion 312 includes two example engines plotted—example322 and example 324. Examples 322, 324 each correspond to a geared andducted turbofan engine (see, e.g., exemplary geared, ducted, turbofanengine 544 described below with reference to FIG. 9 , having a gearbox546). More specifically, each of these engines is a geared engine andfurther is a ducted engine (i.e., includes an outer nacelle surroundingthe primary fan). The primary fan specific thrust rating T_(P) for thesecond subregion 312 ranges from 0.23 to 0.35 and the secondary fanspecific thrust rating T_(S) for the second subregion 312 ranges from0.4 to 0.5.

It will be appreciated that the primary fan specific thrust ratingsT_(P) of the engines in second subregion 312 are higher than the enginesin the first subregion 310. Such may generally be a result of theengines being ducted engines, as the fans may be smaller, requiring themto rotate more quickly to generate a desired amount of thrust. Inaddition, the engines plotted as examples 322, 324 may have lower gearratios than the engines plotted as examples 318, 320. While such wouldgenerally result in a reduced secondary fan specific thrust ratingT_(S), as the engines plotted as examples 322, 324 have higher primaryfan specific thrust rating T_(P), such may drive the secondary fanspecific thrust rating T_(S) higher.

Referring still to FIG. 4 , the third and fourth subregions 314, 316include example engines plotted that are all direct drive engines (i.e.,do not include a reduction gearbox between a driving shaft and a fanshaft). Generally for each of these engines, the secondary fan specificthrust rating T_(S) will be lower for a given primary fan specificthrust rating T_(P), as the secondary fan is rotating at the same speedas the primary fan.

In particular, referring first to the third subregion 314, the thirdsubregion 314 includes examples 326 and 328. The engines plotted asexamples 326 and 328 are direct drive, ducted turbofan engines (see,e.g., exemplary direct drive turbofan engine 538 described below withreference to FIG. 8 ). In particular, these engines may define arelatively high bypass ratio for subsonic flight operations. The primaryfan specific thrust rating T_(P) for the third subregion 314 ranges from0.23 to 0.38 and the secondary fan specific thrust rating T_(S) for thethird subregion 314 ranges from 0.26 to 0.4.

By contrast, referring now to the fourth subregion 316, the fourthsubregion 316 includes examples 330 and 332. The engines plotted asexamples 330 and 332 are also direct drive, ducted turbofan engines, butinstead may define relatively low bypass ratios, e.g., for high flightspeed operations favoring smaller fan diameters. In such a manner, ascompared to the engines plotted as examples 326, 328, the enginesplotted as examples 330, 332 may have smaller fans that rotate morequickly, favoring higher speed and lower drag over efficiency. Theprimary fan specific thrust rating T_(P) for the fourth subregion 316ranges from 0.35 to 0.5 and the secondary fan specific thrust ratingT_(S) for the fourth subregion 316 ranges from 0.37 to 0.5.

Further, values for various other parameters of the influencingcharacteristics of an engine defined by Expression (1) are set forthbelow in TABLE 3:

TABLE 1 Ranges appropriate for using Symbol Description Expression (1)R₁/R₃ Tip radius ratio 1.35 to 10, such as 2 to 7, such as 3 to 5, suchas at least 3.5, such as 3.7, such as at least 4, such as up to 10, suchas up to 7 RqR_(Sec.-Fan) Secondary fan radius 0.2 to 0.9, such as 0.2to 0.7, ratio such as 0.57 to 0.67 RqR_(Prim.-Fan) Primary fan radius0.2 to 0.4, such as 0.25 to 0.35 ratio EFP Effective fan 0.15 to 33,such as 2 to 20, parameter such as 2 to 4.5, such as 2.5 to 10, such as4 to 8, such as 3 to 5 V_(C) Prim.- Corrected primary fan 500 feet persecond (fps) to Fan tip speed 2,000 fps, such as 600 fps to 1,800 fpsV_(C) Sec.- Corrected secondary 500 feet per second (fps) to Fan fan tipspeed 2,000 fps, such as 750 fps to 1,900 fps, such as 1,200 fps to1,800 fps T_(P) Primary fan specific 0.08 to 0.59, such as 0.1 to 0.5thrust rating T_(S) Secondary fan specific 0.21 to 0.6, such as 0.26 to0.5 thrust rating EBA Effective Bypass Area 0.2% to 15%, such as 2% to10%

Referring now to FIGS. 5A through 5F and 6A through 6C, exemplary gasturbine engines are illustrated in accordance with one or more exemplaryembodiments of the present disclosure, showing the relationships betweenthe various parameters of Expression (1). In particular, FIGS. 5Athrough 5F provides a table including numerical values corresponding toseveral of the plotted gas turbine engines in FIGS. 5A through 5C. FIGS.6A through 6C are plots of gas turbine engines in accordance with one ormore exemplary embodiments of the present disclosure, showing therelationships between the tip radius ratio (R₁ to R₃; Y-Axis) and theEBA, effective bypass area (X-axis). FIG. 6A highlights a subrange basedon an EFP value subrange. FIG. 6B highlights a subrange for unductedengines. FIG. 6C highlights a subrange for ducted engines. As will beappreciated at least from Expression (1) and the discussion herein, theEBA multiplied by the ratio of primary fan specific thrust rating TP tosecondary fan specific thrust rating TS (see FIG. 4 ) may generallycorrelate to a percentage of thrust provided by a third stream of anengine. In such a manner, it will be appreciated that the subranges ofFIG. 6B (e.g., 406, 408) may be combined with the subranges insubregions 310 and 312 in FIG. 4 , and similarly the subranges of FIG.6C (e.g., 410, 412) may be combined with the subranges in subregions 314and 316 in FIG. 4 . Further, the subrange of area 308 of FIG. 4 , thesubrange in FIG. 6A, or both may be combined with any of the othersubranges herein (e.g., of FIG. 6B or 6C).

Referring particularly to FIG. 6A, a first range 402 and a second range404 are provided. The first range 402 may correspond to an EFP greaterthan or equal to 0.15 and less than or equal to 33. This range, incombination with the broader ranges for primary and secondary specificthrust ratings T_(P), T_(S) and EBA listed in TABLE 1 may correlate to apercent thrust through the third stream % Fn_(3S) between 2% and 50%.Such may result in an engine having a desired propulsive efficiency.

The second range 404 may correspond to an EFP greater than or equal to 2and less than or equal to 20. This range, in combination with thebroader ranges for the primary and secondary specific thrust ratingsT_(P), T_(S) and EBA listed in TABLE 1 may correlate to a narrower rangeof percent thrust through the third stream % Fn_(3S), e.g., between 5%and about 20%. Such may result in an engine having a more preferredpropulsive efficiency.

Referring now to FIG. 6B, a third range 406 is provided along with afourth range 408, which is a subrange of the third range 406. The thirdand fourth ranges 406, 408 relate to unducted engines, such as aturboprop engine (see FIG. 7 ) or an open rotor engine (see FIGS. 1 and2 ). The third range 406 may correspond to an EBA between 0.8% and 6.5%and an R₁ to R₃ value of 3.4 to 6.5. The fourth range 408 may correspondto an EBA between 0.9% and 2.9% and an R₁ to R₃ value of 3.5 to 6.4.

As discussed above, the unducted engines may generally have larger fans,and as a result may have a larger R₁ to R₃ value. Similarly, because ofthe larger fans, the EBA may generally be lower, as the bypass area ofthe primary fan is so large. Notably, the turboprop engines maygenerally have a higher R₁ to R₃ value than the open rotor turbofanengines. For example, the R₁ to R₃ value for turboprop engines may bebetween 4.5 and 6.5, and the R₁ to R₃ value for open rotor engines maybe between 3.5 and 5.

Referring now to FIG. 6C, a fifth range 410 is provided along with asixth range 412, which is a subrange of the fifth range 410. The fifthand sixth ranges 410, 412 relate to ducted engines, such as a geared,ducted turbofan engine (see FIG. 9 ) or a direct drive, ducted, turbofanengine (see FIG. 8 ). The fifth range 410 may correspond to an EBAbetween 1.25% and 8.9% and an R₁ to R₃ value of 1.7 to 4. The sixthrange 412 may correspond to an EBA between 1.35% and 3.8% and an R₁ toR₃ value of 1.7 to 3.9.

As discussed above, the ducted engines may generally have smaller fans,and as a result may have a small R₁ to R₃ value. As a result of thesmaller fans, the EBA may generally be higher, as the bypass area of theprimary fan is not as large as compared to a potential bypass area ofthe secondary fan. Notably, the geared turbofan engines may generallyhave a higher R₁ to R₃ value than the direct drive turbofan engines, asthe gearbox allows for the primary fan to rotate slower than thesecondary fan, allowing for a larger fan without a significant increasein a pressure ratio of the primary fan (and thus without a significantreduction in efficiency). For example, the R₁ to R₃ value for gearedturbofan engines may be between 2.3 and 4, and the R₁ to R₃ value fordirect drive turbofan engines may be between 1.7 and 2.8.

It will be appreciated that although the discussion above is generallyrelating to the open rotor engine 10 described above with reference toFIGS. 1 and 2 , in various embodiments of the present disclosure, therelationships outlined above with respect to, e.g., Expression (1) maybe applied to any other suitable engine architecture. For example,reference will now be made to FIGS. 7 through 9 , each depictingschematically an engine architecture associated with the presentdisclosure.

Each of the gas turbine engines of FIGS. 7 through 9 generally include arotor 502 rotatable about a rotor axis 504 and a turbomachine 506rotatable about a longitudinal axis 508. The rotor 502 corresponds tothe “primary fan” described herein. The turbomachine 506 is surroundedat least in part by a core cowl 510 and includes a compressor section512, a combustion section 514, and a turbine section 516 in serial floworder. In addition to the rotor 502, the gas turbine engines of FIGS. 7through 9 each also include a ducted mid-fan or secondary fan 518. Thegas turbine engines each include a fan cowl 520 surrounding thesecondary fan 518.

Referring still to the gas turbine engines of FIGS. 7 through 9 , thegas turbine engines each also define a bypass passage 522 downstream ofthe respective rotor 502 and over the respective fan cowl 520 and corecowl 510, and further define a third stream 524 extending from alocation downstream of the respective secondary fan 518 to therespective bypass passage 522 (at least in the embodiments depicted; inother embodiments, the third stream 524 may instead extend to a locationdownstream of the bypass passage 522).

Referring particularly to FIG. 7 , the exemplary gas turbine enginedepicted is configured as a turboprop engine 526. In such a manner, therotor 502 (or primary fan) is configured as a propeller, defining arelatively large diameter. Further, the turboprop engine 526 includes anengine shaft 528 driven by the turbomachine 506, a fan shaft 530rotatable with the rotor 502, and a gearbox 532 mechanically couplingthe engine shaft 528 with the fan shaft 530. The gearbox 532 is anoffset gearbox such that the rotor axis 504 is radially offset from thelongitudinal axis 508 of the turboprop engine 526.

Notably, in other embodiments of the present disclosure, a turbopropengine may be provided with a reverse flow combustor.

Referring to FIGS. 8 and 9 , the gas turbine engines are each configuredas turbofan engines, and more specifically as ducted turbofan engines.In such a manner, the gas turbine engines each include an outer nacelle534 surrounding the rotor 502, and the rotor 502 (or primary fan) ofeach is therefore configured as a ducted fan. Further, each of the gasturbine engines includes outlet guide vanes 536 extending through thebypass passage 522 to the outer nacelle 534 from the fan cowl 520, thecore cowl 510, or both.

More specifically, still, the gas turbine engine of FIG. 8 is configuredas a direct drive, ducted, turbofan engine 538. In particular, thedirect drive, ducted, turbofan engine 538 includes an engine shaft 540driven by the turbine section 516 and a fan shaft 542 rotatable with therotor 502. The fan shaft 542 is configured to rotate directly with(i.e., at the same speed as) the engine shaft 540.

By contrast, the gas turbine engine of FIG. 9 is configured as a geared,ducted, turbofan engine 544. In particular, the geared, ducted, turbofanengine 544 includes the engine shaft 540 driven by the turbine section516 and the fan shaft 542 rotatable with the rotor 502. However, theexemplary geared, ducted, turbofan engine 544 further includes a gearbox546 mechanically coupling the engine shaft 540 to the fan shaft 542. Thegearbox 546 allows the rotor 502 to rotate at a slower speed than theengine shaft 540, and thus at a slower speed than the secondary fan 518.

Notably, the exemplary geared, ducted, turbofan engine 544 of FIG. 9further includes a pitch change mechanism 548 operable with the rotor502 to change a pitch of the rotor blades of the rotor 502. Such mayallow for an increased efficiency of the gas turbine engine.

As will be appreciated from the description herein, various embodimentsof a gas turbine engine are provided. Certain of these embodiments maybe an unducted, single rotor gas turbine engine (see FIGS. 1 and 2 ), aturboprop engine (see FIG. 7 ), or a ducted turbofan engine (see FIGS. 8and 9 ). Another example of a ducted turbofan engine can be found inU.S. patent application Ser. No. 16/811,368 (Published as U.S. PatentApplication Publication No. 2021/0108597), filed Mar. 6, 2020 (FIG. 10 ,Paragraph [0062], et al.; including an annular fan case 13 surroundingthe airfoil blades 21 of rotating element 20 and surrounding vanes 31 ofstationary element 30; and including a third stream/fan duct 73 (shownin FIG. 10 , described extensively throughout the application)). Variousadditional aspects of one or more of these embodiments are discussedbelow. These exemplary aspects may be combined with one or more of theexemplary gas turbine engine(s) discussed above with respect to thefigures.

For example, in some embodiments of the present disclosure, the enginemay include a heat exchanger located in an annular duct, such as in athird stream. The heat exchanger may extend substantially continuouslyin a circumferential direction of the gas turbine engine (e.g., at least300 degrees, such as at least 330 degrees).

In one or more of these embodiments, a threshold power or disk loadingfor a fan (e.g., an unducted single rotor or primary forward fan) mayrange from 25 horsepower per square foot (hp/ft²) or greater at cruisealtitude during a cruise operating mode. In particular embodiments ofthe engine, structures and methods provided herein generate powerloading between 80 hp/ft² and 160 hp/ft² or higher at cruise altitudeduring a cruise operating mode, depending on whether the engine is anopen rotor or ducted engine.

In various embodiments, an engine of the present disclosure is appliedto a vehicle with a cruise altitude up to approximately 65,000 ft. Incertain embodiments, cruise altitude is between approximately 28,000 ftand approximately 45,000 ft. In still certain embodiments, cruisealtitude is expressed in flight levels based on a standard air pressureat sea level, in which a cruise flight condition is between FL280 andFL650. In another embodiment, cruise flight condition is between FL280and FL450. In still certain embodiments, cruise altitude is definedbased at least on a barometric pressure, in which cruise altitude isbetween approximately 4.85 psia and approximately 0.82 psia based on asea level pressure of approximately 14.70 psia and sea level temperatureat approximately 59 degrees Fahrenheit. In another embodiment, cruisealtitude is between approximately 4.85 psia and approximately 2.14 psia.It should be appreciated that in certain embodiments, the ranges ofcruise altitude defined by pressure may be adjusted based on a differentreference sea level pressure and/or sea level temperature.

As such, it will be appreciated that an engine of such a configurationmay be configured to generate at least 25,000 pounds and less than80,000 of thrust during operation at a rated speed, such as between25,000 and 50,000 pounds of thrust during operation at a rated speed,such as between 25,000 and 40,000 pounds of thrust during operation at arated speed. Alternatively, in other exemplary aspects, an engine of thepresent disclosure may be configured to generate much less power, suchas at least 2,000 pounds of thrust during operation at a rated speed.

In various exemplary embodiments, the fan (or rotor) may include twelve(12) fan blades. From a loading standpoint, such a blade count may allowa span of each blade to be reduced such that the overall diameter of theprimary fan may also be reduced (e.g., to twelve feet in one exemplaryembodiment). That said, in other embodiments, the fan may have anysuitable blade count and any suitable diameter. In certain suitableembodiments, the fan includes at least eight (8) blades. In anothersuitable embodiment, the fan may have at least twelve (12) blades. Inyet another suitable embodiment, the fan may have at least fifteen (15)blades. In yet another suitable embodiment, the fan may have at leasteighteen (18) blades. In one or more of these embodiments, the fanincludes twenty-six (26) or fewer blades, such as twenty (20) or fewerblades. Alternatively, in certain suitable embodiments, the fan may onlyinclude at least four (4) blades, such as with a fan of a turbopropengine.

Further, in certain exemplary embodiments, the rotor assembly may definea rotor diameter (or fan diameter) of at least 10 feet, such as at least11 feet, such as at least 12 feet, such as at least 13 feet, such as atleast 15 feet, such as at least 17 feet, such as up to 28 feet, such asup to 26 feet, such as up to 24 feet, such as up to 18 feet.

In various embodiments, it will be appreciated that the engine includesa ratio of a quantity of vanes to a quantity of blades that could beless than, equal to, or greater than 1:1. For example, in particularembodiments, the engine includes twelve (12) fan blades and ten (10)vanes. In other embodiments, the vane assembly includes a greaterquantity of vanes to fan blades. For example, in particular embodiments,the engine includes ten (10) fan blades and twenty-three (23) vanes. Forexample, in certain embodiments, the engine may include a ratio of aquantity of vanes to a quantity of blades between 1:2 and 5:2. The ratiomay be tuned based on a variety of factors including a size of the vanesto ensure a desired amount of swirl is removed for an airflow from theprimary fan.

Additionally, in certain exemplary embodiments, where the engineincludes the third stream and a mid-fan (a ducted fan aft of theprimary, forward fan), a ratio R1/R2 may be between 1 and 10, or 2 and7, or at least 3.3, at least 3.5, at least 4 and less than or equal to7, where R1 is the radius of the primary fan and R2 is the radius of themid-fan.

It should be appreciated that various embodiments of the engine, such asthe single unducted rotor engine depicted and described herein, mayallow for normal subsonic aircraft cruise altitude operation at or aboveMach 0.5. In certain embodiments, the engine allows for normal aircraftoperation between Mach 0.55 and Mach 0.85 at cruise altitude. In stillparticular embodiments, the engine allows for normal aircraft operationbetween Mach 0.75 and Mach 0.85. In certain embodiments, the engineallows for rotor blade tip speeds at or less than 750 feet per second(fps). In other embodiments, the rotor blade tip speed at a cruiseflight condition can be 650 to 900 fps, or 700 to 800 fps.Alternatively, in certain suitable embodiments, the engine allows fornormal aircraft operation of at least Mach 0.3, such as with turbopropengines.

A fan pressure ratio (FPR) for the primary fan of the fan assembly canbe 1.04 to 2.20, or in some embodiments 1.05 to 1.2, or in someembodiments less than 1.08, as measured across the fan blades of theprimary fan at a cruise flight condition.

In order for the gas turbine engine to operate with a fan having theabove characteristics to define the above FPR, a gear assembly may beprovided to reduce a rotational speed of the fan assembly relative to adriving shaft (such as a low pressure shaft coupled to a low pressureturbine). In some embodiments, a gear ratio of the input rotationalspeed to the output rotational speed is between 3.0 and 4.0, between 3.2and 3.5, or between 3.5 and 4.5. In some embodiments, a gear ratio ofthe input rotational speed to the output rotational speed is greaterthan 4.1. For example, in particular embodiments, the gear ratio iswithin a range of 4.1 to 14.0, within a range of 4.5 to 14.0, or withina range of 6.0 to 14.0. In certain embodiments, the gear ratio is withina range of 4.5 to 12 or within a range of 6.0 to 11.0.

With respect to a turbomachine of the gas turbine engine, thecompressors and/or turbines can include various stage counts. Asdisclosed herein, the stage count includes the number of rotors or bladestages in a particular component (e.g., a compressor or turbine). Forexample, in some embodiments, a low pressure compressor may include 1 to8 stages, a high-pressure compressor may include 4 to 15 stages, ahigh-pressure turbine may include 1 to 2 stages, and/or a low pressureturbine (LPT) may include 1 to 7 stages. In particular, the LPT may have4 stages, or between 4 and 7 stages. For example, in certainembodiments, an engine may include a one stage low pressure compressor,an 11 stage high pressure compressor, a two stage high pressure turbine,and 4 stages, or between 4 and 7 stages for the LPT. As another example,an engine can include a three stage low-pressure compressor, a 10 stagehigh pressure compressor, a two stage high pressure turbine, and a 7stage low pressure turbine.

A core engine is generally encased in an outer casing defining one halfof a core diameter (Dcore), which may be thought of as the maximumextent from a centerline axis (datum for R). In certain embodiments, theengine includes a length (L) from a longitudinally (or axial) forwardend to a longitudinally aft end. In various embodiments, the enginedefines a ratio of L/Dcore that provides for reduced installed drag. Inone embodiment, L/Dcore is at least 2. In another embodiment, L/Dcore isat least 2.5. In some embodiments, the L/Dcore is less than 5, less than4, and less than 3. In various embodiments, it should be appreciatedthat the L/Dcore is for a single unducted rotor engine.

The reduced installed drag may further provide for improved efficiency,such as improved specific fuel consumption. Additionally, oralternatively, the reduced installed drag may provide for cruisealtitude engine and aircraft operation at the above describe Machnumbers at cruise altitude. Still particular embodiments may providesuch benefits with reduced interaction noise between the blade assemblyand the vane assembly and/or decreased overall noise generated by theengine by virtue of structures located in an annular duct of the engine.

Additionally, it should be appreciated that ranges of power loadingand/or rotor blade tip speed may correspond to certain structures, coresizes, thrust outputs, etc., or other structures of the core engine.However, as previously stated, to the extent one or more structuresprovided herein may be known in the art, it should be appreciated thatthe present disclosure may include combinations of structures notpreviously known to combine, at least for reasons based in part onconflicting benefits versus losses, desired modes of operation, or otherforms of teaching away in the art.

Although depicted above as an unshrouded or open rotor engine in theembodiments depicted above, it should be appreciated that aspects of thedisclosure provided herein may be applied to shrouded or ducted engines,partially ducted engines, aft-fan engines, or other gas turbine engineconfigurations, including those for marine, industrial, oraero-propulsion systems. Certain aspects of the disclosure may beapplicable to turbofan, turboprop, or turboshaft engines. However, itshould be appreciated that certain aspects of the disclosure may addressissues that may be particular to unshrouded or open rotor engines, suchas, but not limited to, issues related to gear ratios, fan diameter, fanspeed, length (L) of the engine, maximum diameter of the core engine(Dcore) of the engine, L/Dcore of the engine, desired cruise altitude,and/or desired operating cruise speed, or combinations thereof.

This written description uses examples to disclose the presentdisclosure, including the best mode, and also to enable any personskilled in the art to practice the disclosure, including making andusing any devices or systems and performing any incorporated methods.The patentable scope of the disclosure is defined by the claims, and mayinclude other examples that occur to those skilled in the art. Suchother examples are intended to be within the scope of the claims if theyinclude structural elements that do not differ from the literal languageof the claims, or if they include equivalent structural elements withinsubstantial differences from the literal languages of the claims.

Further aspects are provided by the subject matter of the followingclauses:

A gas turbine engine defining a centerline and a circumferentialdirection, the gas turbine engine comprising: a turbomachine comprisinga compressor section, a combustion section, and a turbine sectionarranged in serial flow order, the turbomachine comprising an inletsplitter defining in part an inlet to a working gas flowpath and a fanduct splitter defining in part an inlet to a fan duct flowpath; aprimary fan driven by the turbomachine defining a primary fan tip radiusR1, a primary fan hub radius R2, and a primary fan specific thrustrating TP; a secondary fan located downstream of the primary fan anddriven by the turbomachine, at least a portion of an airflow from theprimary fan configured to bypass the secondary fan, the secondary fandefining a secondary fan tip radius R3, a secondary fan hub radius R4,and a secondary fan specific thrust rating TS; wherein the gas turbineengine defines an Effective Bypass Area, and wherein a ratio of R1 to R3equals

${\frac{R_{1}}{R_{3}} = \sqrt{\left( {EFP} \right)\frac{\left( {1 - {RqR}_{{Sec}.{- {Fan}}}^{2}} \right)}{\left( {1 - {RqR}_{{Prim}.{- {Fan}}}^{2}} \right)}\left( \frac{T_{P}}{T_{S}} \right)\left( {EBA} \right)}};$

wherein EFP is between 0.15 and 33, wherein RqRPrim.-Fan is a ratio ofR1 to R2, wherein RqRSec.-Fan is a ratio of R3 to R4, wherein theprimary fan specific thrust rating TP is between 0.08 and 0.59, whereinthe secondary fan specific thrust rating TS is between 0.21 and 0.6, andwherein the Effective Bypass Area is between 0.2% and 15%.

The gas turbine engine of any of the preceding clause, wherein the ratioof R1 to R3 is between 1.35 and 10.

The gas turbine engine of any preceding clause, wherein EFP is between 2and 20.

The gas turbine engine of any preceding clause, wherein the EffectiveBypass Area is between 2% and 10%.

The gas turbine engine of any preceding clause, wherein RqRPrim.-Fan isbetween 0.2 and 0.4.

The gas turbine engine of any preceding clause, wherein RqRPrim.-Fan isbetween 0.25 and 0.35.

The gas turbine engine of any preceding clause, wherein RqRSec.-Fan isbetween 0.2 and 0.9.

The gas turbine engine of any preceding clause, wherein RqRSec.-Fan isbetween 0.35 and 0.7.

The gas turbine engine of any preceding clause, wherein the gas turbineengine is an unducted gas turbine engine, wherein the Effective BypassArea is between 0.8% and 6.5%, wherein R1 to R3 is between 3.4 and 6.5,wherein the primary fan specific thrust rating TP is between 0.1 and0.35, wherein the secondary fan specific thrust rating TS is between0.35 and 0.5.

The gas turbine engine of any preceding clause, wherein the EffectiveBypass Area is between 0.9% and 2.9%, and wherein R1 to R3 is between3.5 and 6.4.

The gas turbine engine of any preceding clause, wherein the gas turbineengine is a ducted gas turbine engine, wherein the Effective Bypass Areais between 1.25% and 8.9%, wherein R1 to R3 is between 1.7 and 4,wherein the primary fan specific thrust rating TP is between 0.23 and0.5, wherein the secondary fan specific thrust rating TS is between 0.28and 0.5.

The gas turbine engine of any preceding clause, wherein the EffectiveBypass Area is between 1.35% and 8.9%, and wherein R1 to R3 is between1.7 and 3.9.

The gas turbine engine of any preceding clause, wherein EFP is between 2and 4.5, wherein the primary fan defines a primary fan corrected tipspeed during operation of the gas turbine engine at the rated speedduring standard day operating conditions, wherein the secondary fandefines a secondary fan corrected tip speed during operation of the gasturbine engine at the rated speed during standard day operatingconditions, wherein the primary fan corrected tip speed is between 600feet per second and 1,800 feet per second, and wherein the secondary fancorrected tip speed is between 1,200 feet per second and 1,800 feet persecond.

The gas turbine engine of any preceding clause, wherein the fan ductflowpath defines an outlet, and wherein the gas turbine engine furthercomprises: a variable geometry component associated with the secondaryfan, wherein the variable geometry component is a stage of variableinlet guide vanes located immediately upstream of the secondary fan, avariable exhaust nozzle located at the outlet of the fan duct flowpath,or both.

The gas turbine engine of any preceding clause, wherein the gas turbineengine defines a bypass airflow passage, wherein the primary fan isconfigured to provide a first portion of a primary fan airflow to thebypass airflow passage and a second portion of the primary fan airflowto the secondary fan, and wherein the secondary fan is configured toprovide a first portion of a secondary fan airflow to the fan ductflowpath as the fan duct airflow and a second portion of the secondaryfan airflow to the working gas flowpath.

The gas turbine engine of any preceding clause, further comprising: aheat exchanger positioned in thermal communication with the fan ductflowpath.

The gas turbine engine of any preceding clause, further comprising: anarray of inlet guide vanes located immediately upstream of the secondaryfan.

The gas turbine engine of any preceding clause, further comprising: anarray of outlet guide vanes located immediately downstream of thesecondary fan and upstream of the fan duct.

The gas turbine engine of any preceding clause, further comprising: avariable geometry exhaust nozzle located at an exit of the fan duct.

The gas turbine engine of any preceding clause, further comprising: afan cowl surrounding the secondary fan located downstream of the primaryfan, the fan cowl defining in part an engine inlet located downstream ofthe primary fan; wherein the turbomachine further comprises a core cowlsurrounding at least in part the compressor section, the combustionsection, and the turbine section, and wherein the fan duct is definedbetween the core cowl and the fan cowl.

1. A gas turbine engine defining a centerline and a circumferentialdirection, the gas turbine engine comprising: a turbomachine comprisinga compressor section, a combustion section, and a turbine sectionarranged in serial flow order, the turbomachine comprising an inletsplitter defining in part an inlet to a working gas flowpath and a fanduct splitter defining in part an inlet to a fan duct flowpath; aprimary fan driven by the turbomachine defining a primary fan tip radiusR₁, a primary fan hub radius R₂, and a primary fan specific thrustrating T_(P); and a secondary fan located downstream of the primary fanand driven by the turbomachine, at least a portion of an airflow fromthe primary fan configured to bypass the secondary fan, the secondaryfan defining a secondary fan tip radius R₃, a secondary fan hub radiusR₄, and a secondary fan specific thrust rating T_(S), wherein the gasturbine engine defines an Effective Bypass Area, and wherein a ratio ofR₁ to R₃ equals${\frac{R_{1}}{R_{3}} = \sqrt{\left( {EFP} \right)\frac{\left( {1 - {RqR}_{{Sec}.{- {Fan}}}^{2}} \right)}{\left( {1 - {RqR}_{{Prim}.{- {Fan}}}^{2}} \right)}\left( \frac{T_{P}}{T_{S}} \right)\left( {EBA} \right)}},$ and wherein EFP is between 0.15 and 33, wherein RqR_(Prim.-Fan) is aratio of R₂ to R₁, wherein RqRSec.-Fan is a ratio of R₄ to R₃, whereinthe primary fan specific thrust rating T_(P) is between 0.08 and 0.59,wherein the secondary fan specific thrust rating T_(S) is between 0.21and 0.6, and wherein the Effective Bypass Area is between 0.2% and 15%.2. The gas turbine engine of claim 1, wherein the ratio of R₁ to R₃ isbetween 1.35 and
 10. 3. The gas turbine engine of claim 1, wherein EFPis between 2 and
 20. 4. The gas turbine engine of claim 1, wherein theEffective Bypass Area is between 2% and 10%.
 5. The gas turbine engineof claim 1, wherein RqR_(Prim.-Fan) is between 0.2 and 0.4.
 6. The gasturbine engine of claim 5, wherein RqR_(Prim.-Fan) is between 0.25 and0.35.
 7. The gas turbine engine of claim 1, wherein RqR_(Sec.-Fan) isbetween 0.2 and 0.9.
 8. The gas turbine engine of claim 7, whereinRqR_(Sec.-Fan) is between 0.2 and 0.7.
 9. The gas turbine engine ofclaim 1, wherein the gas turbine engine is an unducted gas turbineengine, wherein the Effective Bypass Area is between 0.8% and 6.5%,wherein R₁ to R₃ is between 3.4 and 6.5, wherein the primary fanspecific thrust rating T_(P) is between 0.1 and 0.35, wherein thesecondary fan specific thrust rating T_(S) is between 0.35 and 0.5. 10.The gas turbine engine of claim 9, wherein the Effective Bypass Area isbetween 0.9% and 2.9%, and wherein R₁ to R₃ is between 3.5 and 6.4. 11.The gas turbine engine of claim 1, wherein the gas turbine engine is aducted gas turbine engine, wherein the Effective Bypass Area is between1.25% and 8.9%, wherein R₁ to R₃ is between 1.7 and 4, wherein theprimary fan specific thrust rating T_(P) is between 0.23 and 0.5,wherein the secondary fan specific thrust rating T_(S) is between 0.28and 0.5.
 12. The gas turbine engine of claim 11, wherein the EffectiveBypass Area is between 1.35% and 8.9%, and wherein R₁ to R₃ is between1.7 and 3.9.
 13. The gas turbine engine of claim 1, wherein EFP isbetween 2 and 4.5, wherein the primary fan defines a primary fancorrected tip speed during operation of the gas turbine engine at therated speed during standard day operating conditions, wherein thesecondary fan defines a secondary fan corrected tip speed duringoperation of the gas turbine engine at the rated speed during standardday operating conditions, wherein the primary fan corrected tip speed isbetween 600 feet per second and 1,800 feet per second, and wherein thesecondary fan corrected tip speed is between 1,200 feet per second and1,800 feet per second.
 14. The gas turbine engine of claim 1, whereinthe fan duct flowpath defines an outlet, and wherein the gas turbineengine further comprises: a variable geometry component associated withthe secondary fan, wherein the variable geometry component is a stage ofvariable inlet guide vanes located immediately upstream of the secondaryfan, a variable exhaust nozzle located at the outlet of the fan ductflowpath, or both.
 15. The gas turbine engine of claim 1, wherein thegas turbine engine defines a bypass airflow passage, wherein the primaryfan is configured to provide a first portion of a primary fan airflow tothe bypass airflow passage and a second portion of the primary fanairflow to the secondary fan, and wherein the secondary fan isconfigured to provide a first portion of a secondary fan airflow to thefan duct flowpath as the fan duct airflow and a second portion of thesecondary fan airflow to the working gas flowpath.
 16. The gas turbineengine of claim 1, further comprising: a heat exchanger positioned inthermal communication with the fan duct flowpath.
 17. The gas turbineengine of claim 1, further comprising: an array of inlet guide vaneslocated immediately upstream of the secondary fan.
 18. The gas turbineengine of claim 17, further comprising: an array of outlet guide vaneslocated immediately downstream of the secondary fan and upstream of thefan duct.
 19. The gas turbine engine of claim 1, further comprising: avariable geometry exhaust nozzle located at an exit of the fan duct. 20.The gas turbine engine of claim 1, further comprising: a fan cowlsurrounding the secondary fan located downstream of the primary fan, thefan cowl defining in part an engine inlet located downstream of theprimary fan; wherein the turbomachine further comprises a core cowlsurrounding at least in part the compressor section, the combustionsection, and the turbine section, and wherein the fan duct is definedbetween the core cowl and the fan cowl.